Heat shield having strain compliant matrix and method of forming same

ABSTRACT

A heat shield is disclosed. The heat shield may have a honeycomb core having a plurality of intersecting wall portions forming a plurality of cells. A strain compliant material may be applied to the wall portions of the honeycomb core prior to filling of the cells with an ablative material. An ablative material may be used that at least substantially fills the cells of the honeycomb core.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a divisional of U.S. patent application No.12/138,676 filed on Jun. 13, 2008, which is now U.S. Pat. No. 8,206,546,issued Jun. 26, 2012. The entire disclosure of the above application isincorporated herein by reference.

FIELD

The present disclosure relates to heat shields, and more particularly toa heat shield incorporating a strain compliant matrix that is used tocoat a component of the heat shield to reduce the development of stresscracks in an ablative material of the heat shield during a curingoperation.

BACKGROUND

The statements in this section merely provide background informationrelated to the present disclosure and may not constitute prior art.

Spacecraft require safe, lightweight, affordable heat shields forprotecting the spacecraft and its occupants during re-entry of thespacecraft into the Earth's atmosphere. Traditionally, the weight of theheat shield has been an important factor. The higher the weight of theheat shield panels used the lower the payload that the spacecraft willbe able to carry. Previously manufactured heat shields have typicallybeen made from mixtures of epoxy-phenolic resins with fillers, fromphenolic resins with fillers, from carbon-carbon composites withbackside insulation, from quartz-phenolic composites with backsideinsulation, or from Phenolic Infiltrated Carbon Ablator (PICA) material.However, existing solutions can often add significant weight to aspacecraft.

The cost of manufacturing previously developed heat shields for aspacecraft has also been an important concern for designers.Traditionally, the high cost of manufacturing heat shields forspacecraft has contributed significantly to the overall cost ofmanufacture for a spacecraft. The manufacturing of previously developedheat shields has also often required complex manufacturing processes.

When manufacturing heat shields using a curable ablative material, insome limited instances there may be a tendency for the ablative materialto crack during the cure process. This can be due to shrinkage of theablative material during a curing process. It may also be due tocoefficient of thermal expansion differences between the ablativematerial and other substrate materials that help to form the heatshield. In the event any small or large cracks develop, then such crackswould need to be repaired subsequent to completion of the curing processand any following cool-down period.

Options for attempting to prevent cracks from developing could involvemaking relief cuts in one of more layers of material of the heat shield.This may reduce or eliminate the possibility of one or more large cracksdeveloping in the ablative material during the curing process. However,such an approach would still require post-processing work to fill andrepair the intentionally made relief cuts.

Another option for attempting to prevent cracks in the ablative materialduring the cure process would be to add chemicals to the ablativematerial that would attempt to limit shrinkage of the ablative materialduring the cure process. Obviously, this requires modification of thechemical structure of the ablative material, and such may not bedesirable because of the risk that the additional chemicals mightinterfere with the overall effectiveness of the ablative material inacting as a heat shield.

Still another option might be to cure the ablative material of the heatshield to a less mature state. However, this option might not enable theablative material to provide the same level of ablative performance.This option may also introduce potential problems relating to “heatsoak” which may occur in situ. By “heat soak” it is meant the thermalexposure that the material would encounter during flight or missionconditions. During a normal mission of a spacecraft, each component ofthe spacecraft requiring a heat shield experiences different thermalenvironments. With the less mature cure option, if the thermalconditioning experienced by the heat shield during a mission prior tothe point at which the ablative heat shield needs to function issufficient to advance the cure of the ablative material, then it maypotentially cause the same cracking that the less mature state was usedto avoid. As such, this option makes the expected thermal environmentthat the components are expected to experience prior to the point atwhich ablation is required a more important consideration.

Still another option might be to cure the ablative material separatelyfrom the other components of the heat shield to which the ablativematerial will ultimately be secured, and then perform a secondarybonding operation to permanently attach the ablative material layer toone of the other material layers of the heat shield. However, secondarybonding would significantly increase the overall complexity of the heatshield by requiring gaps, joints and surface machining of the bondsurface on the ablative material.

SUMMARY

In one aspect the present disclosure relates to a heat shield. The heatshield may comprise a honeycomb core having a plurality of intersectingwall portions forming a plurality of cells. A strain compliant materialmay be applied to the wall portions of the honeycomb core prior tofilling of the cells with an ablative material. An ablative material maybe used that at least substantially fills the cells of the honeycombcore.

In another aspect the present disclosure relates to a heat shield thatmay comprise a honeycomb core. The honeycomb core may have a pluralityof intersecting wall portions forming a plurality of cells. An ablativematerial may be applied as an ablative material sheet through anapplication of force to an outer surface edge of the honeycomb core.This may divide the ablative material sheet in a manner to at leastsubstantially fill the cells of the honeycomb core as portions of theablative material sheet are forced into the cells of the honeycomb core.

In still another aspect the present disclosure may relate to heatshield. The heat shield may have a core having a plurality ofintersecting wall portions forming a plurality of cells. A straincompliant material may be included which is applied to the wall portionsof the core prior to filling of the cells with an ablative material. Anablative material may also be included which is applied as an ablativematerial sheet through an application of force to an outer surface edgeof the core. This may result in edge portions of the outer surface edgeof the core dividing the ablative material sheet in a manner to at leastsubstantially simultaneously fill the cells of the core as portions ofthe ablative material sheet are forced into the cells of the core.

Further areas of applicability will become apparent from the descriptionprovided herein. It should be understood that the description andspecific examples are intended for purposes of illustration only and arenot intended to limit the scope of the present disclosure.

BRIEF DESCRIPTION OF THE DRAWINGS

The drawings described herein are for illustration purposes only and arenot intended to limit the scope of the present disclosure in any way.

The drawings described herein are for illustration purposes only and arenot intended to limit the scope of the present disclosure in any way.

FIG. 1 is a side view of one exemplary spacecraft making use of a heatshield formed in accordance with the teachings of the presentdisclosure;

FIG. 2 is a perspective view of one section of the heat shield shown inFIG. 1;

FIG. 3 is a cross section of the heat shield in accordance with sectionline 3-3 in FIG. 2;

FIG. 4 shows the carrier panel side of a honeycomb panel that has hadventing slots cut into the cell walls;

FIG. 5 is a flowchart of exemplary operations that may be performed tomake the ablative material that is used in constructing the heat shieldof FIG. 1;

FIG. 6 is an illustration of a mold tool being filled with the ablativematerial;

FIG. 7 shows a rubber caul sheet being placed over the BPA mix and avacuum bag being secured over the filled mold tool of FIG. 5;

FIG. 8 shows the ablative material being debulked prior to being frozen;

FIG. 9 shows the vacuum bagging material being removed from the moldtool and the frozen preform;

FIGS. 10A and 10B show perspective views of the resulting ablativepreform ready to be placed into the freezer for storage or to be pressedinto a honeycomb core;

FIG. 11 is a flowchart illustrating exemplary operations in forming theheat shield of the present disclosure;

FIG. 12 is a partial side cross sectional view of the ablative preformpositioned over the honeycomb core, and with the preform/honeycomb coreassembly positioned within a mold tool that is enclosed within a vacuumbag ready for placing into the autoclave;

FIG. 13 is an exemplary graph of the pressure and heat profiles usedduring green state curing of the assembly shown in FIG. 12;

FIG. 14 is an exemplary graph of the pressure and heat profiles usedduring post curing of the assembly shown in FIG. 12;

FIG. 15 is a perspective view of an exemplary closeout component thatmay be secured to a perimeter edge of the heat shield to close it; and

FIG. 16 is a simplified side view showing how a monolithic heat shieldmay be formed using a plurality of sections of the heat shield describedin the present disclosure.

FIGS. 17A-17E illustrate a sequence of operations showing the overallapproach for a monolithic heat shield that is formed by first attachingthe honeycomb core to a spacecraft structure and then processing theassembly;

FIGS. 18A-18F illustrate a sequence of operations for an alternateapproach for constructing a monolithic heat shield that is formed byfilling the honeycomb core on a tool that matches the heat shieldstructure, then processing, machining and attaching the ablative panelto the spacecraft in one piece; and

FIG. 19 illustrates the honeycomb core being coated with a straincompliant material prior to the frozen preform being applied to thehoneycomb core.

DETAILED DESCRIPTION

The following description is merely exemplary in nature and is notintended to limit the present disclosure, application, or uses.

Referring to FIG. 1, an exemplary spacecraft 10 is shown incorporating aheat shield 12 in accordance with one embodiment of the presentdisclosure. The heat shield 12 protects the spacecraft 10 and itsoccupants from the heat generated during reentry into the Earth'satmosphere, or during planetary entry. While the heat shield 12 is shownon a manned spacecraft, it will be appreciated that the heat shield 12is well suited for use on a wide variety of other manned and unmannedspace vehicles that are expected to encounter high temperatures on theirexterior surfaces during travel through the Earth's, or a planetaryatmosphere. The heat shield 12 is also potentially usable on other formsof vehicles, and possibly even on fixed (i.e., non-mobile) structures.The heat shield 12 may find use on virtually any form of mobile airborneplatform or ground based vehicle, or possibly even on marine vehicles.

Referring to FIGS. 2 and 3, the heat shield 12 is shown in greaterdetail. The heat shield 12 includes a core 14, which in this example isa honeycomb core. For convenience the core 14 will be referred tothroughout the following discussion as the “honeycomb core 14”. Thehoneycomb core has a plurality of intersecting wall portions 15 thatform a plurality of cells 15 a. An ablative material 16 is press fitinto the cells 15 a of the honeycomb core 14. In FIG. 3, the honeycombcore 14 may be secured via an adhesive layer 18 to a carrier structure20.

The honeycomb core 14 may be formed from a fabric of well knownfiberglass, for example Style 120 (E-glass), which is impregnated with aphenolic resin. The honeycomb core 14 may also be formed from a Lenoweave fiberglass or carbon fiber fabric having an open weaveconstruction. This enables the ablative material 16, when compressedinto the cells 15 a of the honeycomb core 14, to fill the cells 15 a andbecome an integral portion of the wall structure of the honeycomb core14. Prior to filling the cells 15 a of the honeycomb core 14 with theablative material 16, the honeycomb core 14 may be cleaned with a radiofrequency (RF) generated plasma field so that its surfaces arethoroughly conditioned for the remaining manufacturing operations towhich the honeycomb core 14 will be subjected. The plasma field cleaningtreatment is a process that is commercially available. One such companyperforming this process is 4^(th) State, Inc., of Belmont, Calif.

Referring to FIG. 4, following cleaning, and prior to filling the cells15 a with the ablative material 16, the walls 15 of the honeycomb core14 are partially slotted, preferably using a diamond edged cutting tool,on the side of the honeycomb core 14 that will be bonded to the heatshield carrier structure 20. In FIG. 4 these slots are identified byreference numeral 21 and shown in detail on a piece of the honeycombcore 14. The slots 21 provide escape paths for ambient air that mightotherwise create back pressure in the cells 15 a during the subsequenthoneycomb core 14 filling process, and for water and gases that evolveduring a subsequently performed autoclave curing process. The air, waterand gases are drawn off by a vacuum that is applied to a vacuum bagenclosing a preform that forms the ablative material 16, the honeycombcore 14, the carrier structure 20 and the tool. This process will bedescribed in greater detail in the following paragraphs.

The carrier structure 20 may be formed as a multilayer structure fromone or more metal sheets, or possibly even as a honeycomb structurehaving metal, for example titanium, face sheets. For convenience thecarrier structure 20 has been drawn as a single metal layer in FIG. 3.The adhesive layer 18 may be formed by any suitable adhesive, but in oneexample HT-424 adhesive, which is an epoxy-phenolic structural filmadhesive commercially available from Cytec Industries, Inc. of WestPaterson, N.J., is used as the adhesive.

The ablative material 16 is uniquely formulated to form a lightweight,medium density, syntactic foam ablator compound. The ablative material16 may comprise a mixture of phenolic resin, carbon fibers, silica (SiO₂or Manville “Q”) fibers, phenolic microspheres and silica microspheres.The silica fibers and carbon fibers are used to provide structuralreinforcement to the ablative material 16 and to enhance the thermal andablative performance. The silica microspheres and phenolic microspheresare used as density reducing fillers that also enhance the thermal andablative performance of the ablative material 16. The phenolic resin, inone example, may be Plenco 11956 phenolic resin. The silica fibers mayhave a diameter of about 1.5 um. The carbon fibers may be milled fibershaving a length of about 150 um and a diameter of about 7.4 um. Onespecific carbon fiber that is suitable for use is Asbury Graphite MillsAGM-94 milled carbon fibers. The silica microspheres may have a diameterof between about 20-250 um; and the phenolic microspheres may have amost common diameter of between about 20-100 um. In one implementationPlenocet BJO-0930 phenolic microspheres are used. It will beappreciated, however, that all of these dimensions may be varied to suitthe needs of a particular application.

The use of Plenco 11956 resin is particularly advantageous because it isa single component, water based resole phenolic resin that does notrequire adding flammable solvents, toxic curing agents, or reactivediluents to the basic phenolic resin, as with epoxy-novolac or someother types of phenolic resins. Because it is a liquid at roomtemperature it does not have to be heated to be blended with fillers.Because pure phenolic resin is a better ablator than typical curingagents or reactive diluents, the absence of such curing agents andreactive diluents from the phenolic resin helps to provide the ablativematerial 16 with superior thermodynamic response characteristics. Thelack of a curing agent also allows the freshly mixed ablative material16 to have a longer room temperature working life, since rapid cure doesnot initiate until it is heated to above 15° F. The fresh resin has arelatively long storage life at 0° F. of typically about four months andthe room temperature working life for the ablative material 16 is fivedays, unlike epoxy-novolac, or some other phenolic resin based ablators.These characteristics of Plenco 11956 resin, along with the use offrozen preforms, give the ablative material 16 the working time neededto apply it to large monolithic structures that can be cured in onepiece. Some other phenolic resins or epoxy-novolac material systemseither have short working lives that limit the area of the heat shieldthat can be processed at one time, or they require that the ablativematerial be hand injected into individual honeycomb cells using heatedcaulking guns.

Referring to flowchart 50 of FIG. 5, in forming the ablative material16, it is preferred, but not absolutely essential, that the phenolicmicrospheres are dried out using a heated inert atmosphere before theyare used to make the ablative material 16, as indicated at operation 52.This serves to remove any water and volatiles that may remain in themicrospheres from the manufacturing process used to make them, and itstabilizes the state of cure of the phenolic resin that comprises themicrospheres. The silica microspheres may also be dried by the sameprocess to remove adsorbed water. It is also preferred that the variousconstituent materials used to form the ablative material 16 are added ina specific order to avoid clumping, achieve even wetting of the fibersand the microspheres by the phenolic resin, and to obtain uniformblending of all ingredients. The mixing may be done in a commercialbread mixer that imparts high shear forces to the blend but does notchop or mill the fibers and microspheres. To this end, at operation 54 aquantity of phenolic resin is provided, which as explained above ispreferably Plenco 11956 phenolic resin. At operation 56 the silicafibers are added to the resin and mixed to achieve uniform dispersionand wetting by the resin. At operation 58 the carbon fibers are thenadded to the phenolic resin/silica fiber blend and the mixing iscontinued. At operation 60 the silica microspheres are added to theblend and the mixing is continued. At operation 62 the phenolicmicrospheres are added to the blend. At operation 64 the mixing iscontinued until the final uniform wetting and consistency are achieved.The sequence of adding ingredients and mixing is carried out over a timespan of typically between about 23 minutes-30 minutes. In laboratorytesting the ablative material 16 had a density (virgin) of 0.577 g/cm³(36 lbm/ft³); a thermal conductivity (virgin) at room temperature of0.107 W/m-° K (0.62 Btu/hr-ft²); an ablation onset temperature, inNitrogen, of 396° C. (744° F.); a tensile strength through its thicknessof 4.08 MPa (592 lb/in²) and an effective heat of ablation of 69.9×10³KJ/Kg (at a plasma arc jet heat flux of 420 W/cm²). At this point theablative material 16 is ready to be manufactured into a preform.

Referring to FIGS. 6-10, one method for forming a preform comprised ofthe ablative material 16 will be described. In FIG. 6 the ablativematerial 16 is used to fill a mold 70. The mold will have dimensions oflength, width and thickness that correspond to the desired dimensionsfor the preform. In FIG. 7 mold 70 is vacuum bagged with suitablebagging materials 72 and the ablative material 16 is debulked, (i.e.,compressed to consolidate the granules of ablative material and removevoids by means of a partial vacuum drawn on the vacuum bag). In FIG. 8the mold 70 with the ablative material 16 is frozen at approximately−10° F. for about 4-6 hours. In FIG. 9 the bagging material 72 isremoved from the mold. The frozen ablative material preform 74 is shownin FIGS. 10A and 10B. The preforms can be used immediately for fillinghoneycomb core 14, or they can be freezer stored up to 2 months forlater use. For a large heat shield 12, that requires a plurality ofpreforms, the necessary quantity of preforms are made in advance andfreezer stored until the time of final assembly.

Referring now to the flowchart 100 of FIG. 11 and the drawing of FIG.12, a description of using the preform 74 to form the heat shield 12will be described. At operation 102 the plasma cleaned honeycomb core 14is slotted on the carrier structure side (as shown in FIG. 4), using adiamond edged cutting tool, to thus form the slots 21. At operation 104the carrier structure 20 is bonded to the honeycomb core 14. Atoperation 106 the honeycomb core 14 with the carrier structure 20 bondedthereto is placed in a mold tool 75 sized approximately to the dimensionof the honeycomb core and its carrier structure. At operation 108 thefrozen preform 74 is placed over an upper surface of the honeycomb core14, that being the surface opposite to that which the carrier structure20 is secured to. The entire assembly is covered with vacuum baggingmaterials 77 as indicated at operation 110. The assembly ready to beautoclave cured is shown in simplified form in FIG. 12. The core slots21 that provide venting during filling and cure are shown in FIG. 12 atthe intersection of the honeycomb core 14 and the carrier structure 20.

Referring further to FIG. 11, the assembly of FIG. 12 is then autoclavedcured to the “green state”, i.e. partially cured, as indicated atoperation 112. Exemplary temperatures and pressures that may be usedduring the autoclave cure cycle are illustrated in the graph shown inFIG. 13. During the autoclave curing cycle the preform 74 is thawed andsqueezed into the cells 15 a of the honeycomb core 14 under pressureuntil the cells 15 a are completely filled with the material of thepreform 74, to thus form the heat shield 12. Rather than by autoclaveprocessing, the preform 74 could also be compressed into the cells 15 aof the honeycomb core 14 by a mechanical press, as indicated by dashedlines 79.

Once the green state cure operation is complete, the heat shield 12 isremoved from the mold tool 75 (or the mold tool disassembled), asindicated at operation 114. At operation 116 excess material from thepreform 74 that extends above the honeycomb core 14 may be removed bymachining or other means, and the edges, or periphery, of the heatshield may be beveled to reduce the effect of shrinkage stresses duringthe subsequent postcure operation. At operation 118 the heat shield 12is returned to the autoclave for postcure to the final cure state of theablative material 16. Exemplary temperatures and pressures that may beused during the autoclave postcure cycle are illustrated in the graphshown in FIG. 14.

At operation 120 non-destructive examination of the heat shield 12 byx-ray and ultrasonic methods may be performed to: 1) verify theintegrity of the adhesive bonds between the cured ablative material 16,the honeycomb core 14, and the carrier structure 20; 2) verify that thehoneycomb cells are all filled completely, top to bottom, (i.e. the cellfill is free of voids); and 3) verify that there are no internal cracksin the cured ablative material 16. At operation 122 the final outer moldline (OML) contour of the heat shield may be machined, if needed, toprovide a particular, desired contour. At operation 124, edge closeouts125, one of which is shown in FIG. 15, that have been manufactured inthe same manner as the heat shield 12 described above, may be securedsuch as by an adhesive to the carrier structure 20 to close off theexposed peripheral edges of the heat shield 12. The edge closeouts 125may also be adhesively bonded to the edges of the main portion of theheat shield 12, or they may be bonded only to the carrier structure 20and the gaps between the closeouts 125 and the main portion of the heatshield 12 subsequently filled with a room temperature curing siliconeelastomer. Also at operation 124 the edge closeouts 125 of the heatshield 12 are all non-destructively inspected.

Referring to FIG. 16, when making a monolithic non-planar heat shieldfor the spacecraft 10 shown in FIG. 1, a slotted honeycomb core 150,similar or identical in construction to honeycomb core 14, may be formedwith the desired moldline needed to enable attachment of a finished heatshield to the outer surface of the spacecraft. In this regard aplurality of ablative preforms 74 may be cut to desired shapes and layedonto the honeycomb core 150, which has been secured with an adhesivelayer 18 to the carrier structure 20, and then the entire assemblyvacuum bagged and cured in an autoclave as a single piece assembly. Eachof the preforms 74 may have chamfered edges 74 a to help interlock withadjacently placed preforms. An alternative for making a monolithic,non-planar heat shield as shown in FIG. 16 is to temporarily secure theslotted honeycomb core 14 or 150 to a male tool that matches the outermold line contour of the spacecraft structure, hen layup, vacuum bag,press into the core, and autoclave cure the ablative material 16 of thefrozen preform 74 in the manner described above. Then the monolithicablator is removed from the tool, the inner mold line contour isverified or machined, and the single piece is secured by an adhesive tothe spacecraft structure outer mold line.

In one variation of producing the ablative material 16, the material 16may be forced through a mesh screen, for example a 100 mesh screen,(meaning a stainless steel wire screen with 100 openings per inch thatare about 0.005 inches on a side), to form a pelletized ablativematerial. The pelletized ablative material may then be distributed overthe cells of the honeycomb core 14 to completely fill the cells prior tovacuum bagging of the honeycomb core.

The methodology of the present disclosure thus provides a means forfilling large areas of a honeycomb core structure at one time ratherthan filling each cell individually, or by filling tile size pieces ofhoneycomb core by machine or hand pressing material into both sides ofthe core, followed by curing and then machining to a finished shape forinstallation. This approach makes possible at least three major optionsfor heat shield assembly. The first option is highly advantageous andinvolves pre-bonding the unfilled honeycomb core to the exterior of aspacecraft using an existing adhesive that has been certified for mannedspaceflight. Thus, the present disclosure eliminates the need for thedevelopment and certification of a new attachment design for attachingthe heat shield to a spacecraft using gore segments or tiles. The secondadvantageous option that the methodology discussed herein makes possibleis that when a particular spacecraft design does not allow forprocessing the heat shield on the spacecraft, a single piece monolithicablator assembly may be made on the side and then secured to thespacecraft by an adhesive in one operation. The third option, which hasadvantages for some spacecraft as well as for hypersonic aircraft,ground vehicles and stationary applications, is to make large preformedcured billets that are subsequently machined into panels, gores or largeand small tiles. Options 1 and 2 are illustrated and compared in FIGS.17A-17E and 18A-18F, respectively. In FIG. 17A the honeycomb core 14 isfirst bonded to the carrier structure 20 a that will ultimately form aportion of a spacecraft. In FIG. 17B the assembled honeycomb core 14 andcarrier structure 20 are then cured in an autoclave 160. The honeycombcore 14 may then be filled with the ablative material 16 and then curedin the autoclave 160 (FIG. 17C). The outer mold line (OML) of theresulting cured assembly of FIG. 17C may then be machined to the desiredshape and/or contour, as indicated in FIG. 17D. The resulting product isshown in FIG. 17E. In FIGS. 18A-18F, option two described above isillustrated. The ablative material 16 is first compressed into thehoneycomb core 14. In FIG. 18B, the assembly shown in FIG. 18A is thenautoclave cured in the autoclave 160 to form assembly 170. In FIG. 18C amachine tool 180 is used to machine the inner mold line (IML) of thecured assembly 170 to form machined assembly 185, which is shown in FIG.18D. In FIG. 18D the machined assembly 185 is then bonded to the carrierstructure 20 to form assembly 190. In FIG. 18E the outer mold line (OML)of the assembly 190 is machined with a machine tool 200. The finishedproduct 205 is shown in FIG. 18F.

The heat shield 12 manufactured as described above is lighter thanexisting heat shields made from pre-existing approaches because of thegreater mass efficiency of the ablator composition. The heat shield 12also uses safe, non-toxic materials. The heat shield 12 allows twooptions for a monolithic heat shield design to be constructed that canbe made in accordance with less complex manufacturing procedurestraditionally employed in the manufacture of such heat shields. Thesebenefits also help to reduce the cost of the heat shield 12 as well asthe time needed to manufacture it. In particular, curing the ablativematerial 16 after it has been attached to the honeycomb/carrierstructure avoids the need to form or machine a cured ablative materialto match the contour of a heat shield carrier structure, which must takeinto account machining errors and any variations in each specificcarrier structure piece.

Referring to FIG. 19, a system and method for providing a straincompliant coating to the heat shield 12 will be described. In someinstances the curing of the frozen ablative preform 74 (FIGS. 10A and10B) in an autoclave may cause a small or moderate degree of shrinkageof the ablative material 16. Since the honeycomb core 14 may be bondedto the carrier structure 20 before the frozen preform 74 is compressedinto the cells 15 a of the honeycomb core, and since the carrierstructure 20 will typically be comprised of one or more metal layers ofmaterial, when curing occurs the differences in the thermal coefficientsof expansion between the ablative material 16 and the carrier structure20 may occasionally introduce stress cracks into the cured ablativematerial 16. These stress cracks would need to be repaired before themanufacture of the heat shield 12 can be considered complete.

A highly advantageous method for substantially reducing, or entirelyeliminating, the risk of stress cracks developing in the ablativematerial 16 is to coat the surface of the honeycomb core 14 with astrain compliant material before the frozen fiber preform 74 iscompressed into the cells 15 a. This may be accomplished by dipping thehoneycomb core 14, as shown in FIG. 19, into a reservoir 300 that isfilled with a strain compliant material 302 so that the surfaces of thewall portions 15 are completely covered with the strain compliantmaterial. In one embodiment the strain compliant material may comprise atwo-part, chemically cured, silicone plastic material. In one specificembodiment RTV-560 silicone rubber compound is used. In anotherimplementation the strain compliant material 302 may comprise a powderedrubber mixed with a matrix or resin as the binder. The resin or matrixcould be the same resin or matrix that is to form the ablative material16. In another implementation the strain compliant material 302 maycomprise cork in powder form mixed with a binder such as resin ormatrix. In still another implementation the strain compliant materialmay comprise a thermoplastic polymer.

The strain compliant material 302 will typically be in a viscous formthat enables it to flow easily over all of the surfaces of the honeycombcore 14 while the honeycomb core 14 is submerged therein within thereservoir 300. It may also be possible to spray the honeycomb core 14with the strain compliant material 302 via a spray gun provided theconsistency of the precise formulation of the strain compliant materialused 302 permits such a manner of application. One particular straincompliant material that would be suitable for application via a spraygun is nylon. Still further, the strain compliant material could beapplied via a brush. However, it is anticipated that applying the straincompliant material 302 by dipping the entire honeycomb core 14 into abath of the material 302 is likely to be the most expeditious way ofcoating the honeycomb core 14.

The strain compliant material 302 may also have added to it a powder,for example silica microballoons or high purity silica powder, that aidsthe ablative material 16 of the frozen preform 74 in adhering to thewall portions 15 of the honeycomb core 14 during the cure process. Thepowder may help to provide a more slightly textured surface that enablesthe ablative material 16 to better “grip” onto the strain compliantmaterial. Alternative types of powders could comprise zirconia oralumina powder.

The strain compliant material 302 effectively increases the straincompliance of the heat shield 12 without the need to reformulate theablative material 16 itself. This increased strain compliance reducesinternal stresses that result from shrinkage when the frozen preform 74is being compressed into the cells 15 a of the honeycomb core 14 andcured through the application of heat in an autoclave. If the straincompliant material 302 is a thermoplastic, it will soften during thecuring operation when heat is being applied to the honeycomb core 14.The thermoplastic will then “flow” or, deform while the ablativematerial 16 undergoes shrinkage. After the shrinkage has ceased, thehoneycomb core 14 with the ablative material 16 filling its cells 15 amay be cooled, during which time the thermoplastic (i.e., forming thestrain compliant material 302) will again harden. The strain compliantmaterial 302, once re-hardened during the cool down period (i.e.,following the cure operation) will contribute additional strength andstiffness to the completed heat shield 12.

The use of the strain compliant material 12 provides significantmanufacturing benefits. For one, the ablative material 16 does not needto be reformulated in an attempt to control shrinkage during the cureprocess. Another benefit is that the ablative material 16 can be fullycured in a single operation, rather than using a less than mature curecycle. Still another benefit is that there is no need to make reliefcuts in the ablative material 16 to try and control the formation ofstress cracks during the cure cycle. Still another benefit is that thestrain compliant material 302 provides increased strain compliance thatenables the heat shield 12 to better accommodate in-flight stressesresulting from structural deflections or deformations that the heatshield is subjected to.

While various embodiments have been described, those skilled in the artwill recognize modifications or variations which might be made withoutdeparting from the present disclosure. The examples illustrate thevarious embodiments and are not intended to limit the presentdisclosure. Therefore, the description and claims should be interpretedliberally with only such limitation as is necessary in view of thepertinent prior art.

What is claimed is:
 1. A heat shield comprising: a honeycomb core havinga plurality of intersecting wall portions forming a plurality of cells,the honeycomb core having an open weave construction; a strain compliantmaterial coated on the wall portions of the honeycomb core; and anablative material that at least substantially fills the cells of thehoneycomb core, the ablative material including carbon fibers to providestructural reinforcement to the ablative material; wherein: a portion ofthe ablative material is compressed into the open weave construction ofthe honeycomb core and is an integral portion of the wall portions ofthe honeycomb core.
 2. The heat shield of claim 1, wherein the straincompliant material comprises a two-part, chemically cured, Siliconeplastic material.
 3. The heat shield of claim 2, wherein said two-part,chemically cured, silicone plastic material comprises RTV-560 Siliconeplastic.
 4. The heat shield of claim 1, wherein the strain compliantmaterial comprises powdered rubber mixed with a binder.
 5. The heatshield of claim 1, wherein the strain compliant material comprises corkin powder form that is mixed with a binder.
 6. The heat shield of claim1, wherein the strain compliant material comprises a thermoplasticpolymer.
 7. The heat shield of claim 1, wherein the strain compliantmaterial comprises nylon applied to said wall portions while in aviscous form.
 8. The heat shield of claim 1, wherein the straincompliant material further comprises a powder to aid in attachment ofthe ablative material to the strain compliant material duringmanufacture of the heat shield.
 9. The heat shield of claim 1, furthercomprising a carrier structure formed from at least one metallic sheetof material, and wherein the carrier structure is bonded to a surface ofthe honeycomb core.
 10. A heat shield comprising: a honeycomb corehaving a plurality of intersecting wall portions forming a plurality ofcells, selected ones of the wall portions further being slotted along acommon edge of the honeycomb core, the honeycomb core having an openweave construction; and an ablative material, applied as an ablativematerial sheet through an application of force to an outer surface edgeof the honeycomb core, that divides the ablative material sheet in amanner to at least substantially fill the cells of the honeycomb core asportions of the ablative material sheet are forced into the cells of thehoneycomb core, wherein: the slotted selected ones of the wall portionsfacilitate de-gassing of the ablative material during manufacture; and aportion of the ablative material is compressed into the open weaveconstruction of the honeycomb core and is an integral portion of thewall portions of the honeycomb core: and a strain compliant materialcoated on the wall portions of the honeycomb core.
 11. The heat shieldof claim 10, wherein the strain compliant material comprises a two-part,chemically cured, Silicone plastic material.
 12. The heat shield ofclaim 11, wherein the two-part, chemically cured, silicone plasticmaterial comprises RTV-560 Silicone plastic.
 13. The heat shield ofclaim 10, wherein the strain compliant material comprises at least oneof: powdered rubber mixed with a binder; cork in powder form that ismixed with a binder; and a thermoplastic polymer.
 14. The heat shield ofclaim 10, wherein the strain compliant material comprises at least oneof: nylon applied to said wall portions while in a viscous form; and apowder to aid in attachment of the ablative material to the straincompliant material during manufacture of the heat shield.
 15. A heatshield comprising: a core having a plurality of intersecting wallportions forming a plurality of cells, the wall portions having edgesthat are slotted, the core having an open weave construction; a straincompliant material coated on the wall portions of the core; and anablative material including carbon fibers, applied as an ablativematerial sheet through an application of force to an outer surface edgeof the core, such that edge portions of the outer surface divide theablative material sheet in a manner to at least substantiallysimultaneously fill the cells of the core as portions of the ablativematerial sheet are forced into the cells of the core, the carbon fibersadding structural strength, wherein: the slotted wall portions furtherare arranged on a common side of the core and enable improved de-gassingof the ablative material during manufacture of the heat shield; and aportion of the ablative material is compressed into the open weaveconstruction of the core and is an integral portion of the wall portionsof the core.
 16. The heat shield of claim 15, wherein the core comprisesa honeycomb core.
 17. The heat shield of claim 15, wherein the straincompliant material comprises at least one of: a two-part, chemicallycured, Silicone plastic material; and RTV-560 Silicone plastic.
 18. Theheat shield of claim 15, wherein the strain compliant material comprisesat least one of: powdered rubber mixed with a binder; cork in powderform that is mixed with a binder; a thermoplastic polymer; nylon appliedto said wall portions while in a viscous form; and a powder to aid inattachment of the ablative material to the strain compliant materialduring manufacture of the heat shield.
 19. The heat shield of claim 15,further comprising a carrier structure formed from at least one metallicsheet of material, and wherein the carrier structure is bonded to asurface of the core.